Tip turbine engine and corresponding operating method

ABSTRACT

A tip turbine engine has a more efficient core airflow path from the hollow fan blades, through the combustor and to the combustion chamber of a combustor. The turbine engine includes a rotatable fan having a plurality of radially-extending fan blades each defining compressor chambers extending radially therein. A turbine is mounted to the outer periphery of the fan. A diffuser at a radially outer end of each compressor chamber turns core airflow through the compressor chamber toward the combustor. The high velocity, high pressure core airflow from the compressor chambers in the hollow fan blades is diffused before the compressed core airflow enters the combustor, thereby improving the efficiency of the tip turbine engine. Further, the overall diameter of the tip turbine engine is reduced by the arrangement of the diffuser case in a position not directly radially outward of the fan blades.

The present invention relates to a tip turbine engine, and moreparticularly to a novel core airflow path for a tip turbine engine.

An aircraft gas turbine engine of the conventional turbofan typegenerally includes a forward bypass fan and a low pressure compressor, amiddle core engine, and an aft low pressure turbine, all located along acommon longitudinal axis. A high pressure compressor and a high pressureturbine of the core engine are interconnected by a high spool shaft. Thehigh pressure compressor is rotatably driven to compress air enteringthe core engine to a relatively high pressure. This high pressure air isthen mixed with fuel in a combustor, where it is ignited to form a highenergy gas stream. The gas stream flows axially aft to rotatably drivethe high pressure turbine, which rotatably drives the high pressurecompressor via the high spool shaft. The gas stream leaving the highpressure turbine is expanded through the low pressure turbine, whichrotatably drives the bypass fan and low pressure compressor via a lowspool shaft.

Although highly efficient, conventional turbofan engines operate in anaxial flow relationship. The axial flow relationship results in arelatively complicated elongated engine structure of considerable lengthrelative to the engine diameter. This elongated shape may complicate orprevent packaging of the engine into particular applications.

A recent development in gas turbine engines is the tip turbine engine.Tip turbine engines locate an axial compressor forward of a bypass fanwhich includes hollow fan blades that receive airflow from the axialcompressor therethrough such that the hollow fan blades operate as acentrifugal compressor. Compressed core airflow from the hollow fanblades is mixed with fuel in an annular combustor, where it is ignitedto form a high energy gas stream which drives the turbine that isintegrated onto the tips of the hollow bypass fan blades for rotationtherewith as generally disclosed in U.S. Patent Application PublicationNos.: 20030192303; 20030192304; and 20040025490. The tip turbine engineprovides a thrust-to-weight ratio equivalent to or greater thanconventional turbofan engines of the same class, but within a package ofsignificantly shorter length.

In the known tip turbine engines, most of the core airflow flowsradially outwardly from the radial outer ends of the hollow fan bladesinto the combustor, which is mounted about the periphery of the fan. Afuel injector aft of the fan delivers fuel into the combustor where itis ignited. The high-energy gas stream is then directed axially forwardin the combustor, then redirected radially inward and then turned onceagain axially rearward to pass through turbine blades between the fanblades to rotatably drive the fan. This design has some drawbacks.Mounting the combustor about the periphery of the fan increases theoverall diameter of the known tip turbine engine.

Additionally, in the known tip turbine engines, the compressed airflowfrom the hollow fan blades exits directly into the combustor. A lack ofdiffusion between the centrifugal compressor and the combustor causes alarge loss in efficiency.

SUMMARY OF THE INVENTION

A tip turbine engine according to the present invention has a moreefficient core airflow path from the hollow fan blades through to thecombustion chamber of the combustor.

The turbine engine includes a rotatable fan having a plurality ofradially-extending fan blades each defining compressor chambersextending radially therein. A turbine is mounted to the outer peripheryof the fan. A diffuser at a radially outer end of each compressorchamber turns core airflow through the compressor chamber toward anannular combustor disposed in front of the turbine. Thus, in the tipturbine engine according to the present invention, the high velocity,high pressure core airflow from the compressor chambers in the hollowfan blades is diffused before the compressed core airflow enters thecombustor, thereby improving the efficiency of the tip turbine engine.Further, the overall diameter of the tip turbine engine is reduced bythe arrangement of the diffuser case in a position not directly radiallyoutward of the fan blades.

BRIEF DESCRIPTION OF THE DRAWINGS

Other advantages of the present invention can be understood by referenceto the following detailed description when considered in connection withthe accompanying drawings wherein:

FIG. 1 is a partial sectional perspective view of a tip turbine engineaccording to the present invention.

FIG. 2 is a longitudinal sectional view of the tip turbine engine ofFIG. 1 taken along an engine centerline.

FIG. 3 is an enlarged view of the diffuser, combustor and turbine areaof FIG. 2.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

FIG. 1 illustrates a general perspective partial sectional view of a tipturbine engine (TTE) type gas turbine engine 10. The engine 10 includesan outer nacelle 12, a rotationally fixed static outer support structure14 and a rotationally fixed static inner support structure 16. Aplurality of fan inlet guide vanes 18 are mounted between the staticouter support structure 14 and the static inner support structure 16.Each inlet guide vane preferably includes a variable trailing edge 18A.

A nosecone 20 is preferably located along the engine centerline A toimprove airflow into an axial compressor 22, which is mounted about theengine centerline A behind the nosecone 20.

A fan-turbine rotor assembly 24 is mounted for rotation about the enginecenterline A aft of the axial compressor 22. The fan-turbine rotorassembly 24 includes a plurality of hollow fan blades 28 to provideinternal, centrifugal compression of the compressed airflow from theaxial compressor 22 for distribution to an annular combustor 30 locatedwithin the rotationally fixed static outer support structure 14.

A turbine 32 includes a plurality of tip turbine blades 34 (two stagesshown) which rotatably drive the hollow fan blades 28 relative aplurality of tip turbine stators 36 which extend radially inwardly fromthe rotationally fixed static outer support structure 14. The annularcombustor 30 is disposed axially forward of the turbine 32 andcommunicates with the turbine 32.

Referring to FIG. 2, the rotationally fixed static inner supportstructure 16 includes a splitter 40, a static inner support housing 42and a static outer support housing 44 located coaxial to said enginecenterline A.

The axial compressor 22 includes the axial compressor rotor 46, fromwhich a plurality of compressor blades 52 extend radially outwardly, anda fixed compressor case 50. A plurality of compressor vanes 54 extendradially inwardly from the compressor case 50 between stages of thecompressor blades 52. The compressor blades 52 and compressor vanes 54are arranged circumferentially about the axial compressor rotor 46 instages (three stages of compressor blades 52 and compressor vanes 54 areshown in this example). The axial compressor rotor 46 is mounted forrotation upon the static inner support housing 42 through a forwardbearing assembly 68 and an aft bearing assembly 62.

The fan-turbine rotor assembly 24 includes a fan hub 64 that supports aplurality of the hollow fan blades 28. Each fan blade 28 includes aninducer section 66, a hollow fan blade section 72 and a diffuser section74. The inducer section 66 receives airflow from the axial compressor 22generally parallel to the engine centerline A and turns the airflow froman axial airflow direction toward a radial airflow direction. Theairflow is radially communicated through a core airflow passage 80 whichacts as a compressor chamber within the fan blade section 72 where theairflow is centrifugally compressed. From the core airflow passage 80,the airflow is diffused and turned once again by the diffuser section 74toward an axial airflow direction toward the annular combustor 30.Preferably, the airflow is diffused axially forward in the engine 10,however, the airflow may alternatively be communicated in anotherdirection.

All or substantially all of the airflow through the core airflow passage80 is core airflow directed by the diffuser section 74 axially forwardtoward the combustor 30. Minimal amounts of airflow may be directedradially outwardly from the diffuser section 74 through the turbineblades 34 (paths not shown) to cool the tip turbine blades 34. Thiscooling airflow is then discharged through radially outer ends of thetip turbine blades 34 and then into the annular combustor 30. However,at least substantially all of the airflow is core airflow directed bythe diffuser section 74 toward the combustor 30. As used herein, “coreairflow” is airflow that flows to the combustor 30.

A gearbox assembly 90 aft of the fan-turbine rotor assembly 24 providesa speed increase between the fan-turbine rotor assembly 24 and the axialcompressor 22, which in the embodiment shown is at a 3.34 ratio. In theembodiment shown, the gearbox assembly 90 is a planetary gearbox thatprovides co-rotating engagement between the fan-turbine rotor assembly24 and an axial compressor rotor 46. Alternatively, a counter-rotatingplanetary gearbox could be provided. The gearbox assembly 90 is mountedfor rotation between the static inner support housing 42 and the staticouter support housing 44.

The gearbox assembly 90 includes a sun gear 92, which rotates with theaxial compressor 22, and a planet carrier 94, which rotates with thefan-turbine rotor assembly 24 to provide a speed differentialtherebetween. A plurality of planet gears 93 (one shown) are mounted tothe planet carrier 94. The planet gears 93 engage the sun gear 92 and aring gear 95. The gearbox assembly 90 is mounted for rotation betweenthe sun gear 92 and the static outer support housing 44 through aforward bearing 96 and a rear bearing 99. The forward bearing 96 and therear bearing 99 are both tapered roller bearings and both handle radialloads. The forward bearing 96 handles the aft axial load, while the rearbearing 99 handles the forward axial loads. The sun gear 92 isrotationally engaged with the axial compressor rotor 46 at a splinedinterconnection 100 or the like.

It should be noted that the gearbox assembly 90 could utilize othertypes of gear arrangements or other gear ratios and that the gearboxassembly 90 could be located at locations other than aft of the axialcompressor 22. For example, the gearbox assembly 90 could be located atthe front end of the axial compressor 22. Alternatively, the gearboxassembly 90 could provide a speed decrease between the fan-turbine rotorassembly 24 and the axial compressor rotor 46.

The annular combustor 30 and turbine 32 are shown in greater detail inFIG. 3. The annular combustor 30 is located entirely fore of a fan planeP, within which the fan blades 28 rotate. The annular combustor 30includes an annular combustion chamber 112 defined between an annularinner combustion chamber wall 114 and annular outer combustion chamberwall 116. A forward wall 118 at a forward end of the combustion chamber112 has mounted thereto a fuel injector 120, which directs fuel into thecombustion chamber 112. The combustion chamber 112 includes a combustionchamber outlet 122 opposite the forward wall 118. The combustion chamberoutlet 122 is substantially axially aligned with the forward wall 118such that a substantially axial combustion path 124 is defined throughthe combustion chamber 112. The annular inner and outer combustionchamber walls 114, 116 and the forward wall 118 are perforated to permitcore airflow into the combustion chamber 112.

An annular diffuser case 128 substantially encloses the annular innerand outer combustion chamber walls 114, 116 and the forward wall 118. Aninner diffuser case wall 130 defines a core airflow path 132 with theannular inner combustion chamber wall 114. A core airflow path inlet 134is axially aligned (i.e. along an axis parallel to the engine centerlineA (FIG. 1)) with the diffuser section 74 and is substantially radiallyaligned (i.e. along a radius from engine centerline A) with thecombustion chamber outlet 122. The core airflow path inlet 134 leadsinto the combustion chamber 112 through the annular inner and outercombustion chamber walls 114, 116 and the forward wall 118.

In operation, referring to FIG. 2, air enters the axial compressor 22,where it is compressed by the three stages of the compressor blades 52and compressor vanes 54. The compressed air from the axial compressor 22enters the inducer section 66 in a direction generally parallel to theengine centerline A, and is then turned by the inducer section 66radially outwardly through the core airflow passage 80 of the hollow fanblades 28. The airflow is further compressed centrifugally in the hollowfan blades 28 by rotation of the hollow fan blades 28. From the coreairflow passage 80, the airflow is turned and diffused axially forwardin the engine 10 by diffuser section 74 into core airflow path inlet 134of the annular combustor 30, as shown in FIG. 3. This diffusion improvesthe efficiency, by reducing the losses encountered when the compressedcore airflow enters the larger combustion chamber 112. The diffusedcompressed core airflow from the hollow fan blades 28 then flowsradially outwardly and through the annular inner and outer combustionchamber walls 114, 116 and the forward wall 118 to the combustionchamber 112 where it is mixed with fuel and ignited to form ahigh-energy gas stream.

The high-energy gas stream expands and follows the combustion path 124,which is substantially axial all the way from the forward wall 118 ofthe combustion chamber 112 through the combustion chamber outlet 122 andthrough the tip turbine blades 36. The high-energy gas stream rotatablydrives the plurality of tip turbine blades 34 mounted about the outerperiphery of the fan-turbine rotor assembly 24 to drive the fan-turbinerotor assembly 24, which in turn drives the axial compressor 22 via thegearbox assembly 90. Because the combustion path 124 is substantiallyaxial, the efficiency of the combustor 30 is improved over the knowncombustors in tip turbine engines. Additionally, because the combustor30 is located fore of the fan blades 28 and is not located the fan planeP, the tip turbine engine 10 has a smaller diameter than the known tipturbine engines.

The fan-turbine rotor assembly 24 discharges fan bypass air axially aftto merge with the core airflow from the turbine 32 in an exhaust case106. A plurality of exit guide vanes 108 are located between the staticouter support housing 44 and the rotationally fixed static outer supportstructure 14 to guide the combined airflow out of the engine 10 andprovide forward thrust. An exhaust mixer 110 mixes the airflow from theturbine blades 34 with the bypass airflow through the fan blades 28.

In accordance with the provisions of the patent statutes andjurisprudence, exemplary configurations described above are consideredto represent a preferred embodiment of the invention. However, it shouldbe noted that the invention can be practiced otherwise than asspecifically illustrated and described without departing from its spiritor scope.

1. A turbine engine comprising: a fan rotatable about an axis, the fanincluding a plurality of radially-extending fan blades, at least one ofthe fan blades defining a compressor chamber extending radially therein;a combustor partially defining a combustion path extending axiallywithin a combustion chamber through a combustion chamber outlet, thecombustion path extending axially aft of the combustion chamber outlet;and a diffuser at a radially outer end of the compressor chamber, thediffuser turning core airflow from the compressor chamber toward thecombustor, the diffuser disposed radially inward of the combustion pathand wherein the diffuser turns at least substantially all of the coreairflow from the compressor chamber from a radial direction to an axialdirection toward the combustor.
 2. The turbine engine of claim 1 whereinthe fan is rotatable about the axis in a fan plane and wherein thecombustion path is radially outward of the diffuser in the fan plane. 3.A turbine engine comprising: a fan rotatable about an axis in a fanplane, the fan including a plurality of radially-extending fan blades,at least one of the fan blades defining a compressor chamber extendingradially therein, wherein the combustion path is radially outward of thediffuser in the fan plane; a combustor partially defining a combustionpath extending axially within a combustion chamber through a combustionchamber outlet, the combustion path extending axially aft of thecombustion chamber outlet; and a diffuser at a radially outer end of thecompressor chamber, the diffuser turning core airflow from thecompressor chamber toward the combustor, the diffuser disposed radiallyinward of the combustion path; and a turbine at an outer circumferenceof the fan aft of the combustor, the fan rotatably driven by the turbineand the turbine rotatably driven by the combustor, the combustor pathextending axially rearwardly from the combustion chamber through thecombustion chamber outlet and then through the turbine.
 4. The turbineengine of claim 3 wherein the combustor generates a high-energy gasstream to drive the turbine.
 5. The turbine engine of claim 1 whereinthe combustion path extends axially from a forward end of the combustionchamber to the combustion chamber outlet at a rearward end of thecombustion chamber.
 6. A turbine engine comprising: a fan rotatableabout an axis, the fan including a plurality of radially-extending fanblades, at least one of the fan blades defining a compressor chamberextending radially therein; a combustor partially defining a combustionpath extending axially within a combustion chamber through a combustionchamber outlet, the combustion path extending axially aft of thecombustion chamber outlet, wherein the combustion path extends axiallyfrom a forward end of the combustion chamber to the combustion chamberoutlet at a rearward end of the combustion chamber; and a diffuser at aradially outer end of the compressor chamber, the diffuser turning coreairflow from the compressor chamber toward the combustor, the diffuserdisposed radially inward of the combustion path a turbine at an outercircumference of the fan aft of the combustor, the fan rotatably drivenby the turbine and the turbine rotatably driven by the combustor, thecombustor path extending axially rearwardly from the forward end of thecombustion chamber through the combustion chamber outlet and thenthrough the turbine.
 7. The turbine engine of claim 1 wherein thediffuser turns at least substantially all of the core airflow forward ofthe at least one fan blade.
 8. The turbine engine of claim 1 wherein thecore airflow from the diffuser flows generally axially and then radiallyinto the combustion chamber.
 9. A turbine engine comprising: a fanrotatable about an axis, the fan including a plurality of fan blades, atleast one fan blade defining a compressor chamber therein, thecompressor chamber centrifugally compressing core airflow therein; adiffuser turning at least substantially all of the core airflow from thecompressor chamber from a substantially radial direction toward agenerally axial direction; a combustor receiving core airflow from thediffuser; and a turbine mounted to outer ends of the fan blades radiallyoutwardly of the diffuser and aft of the combustor.
 10. The turbineengine of claim 9 wherein the diffuser turns the core airflow from thecompressor chamber toward an axially forward direction.
 11. The turbineengine of claim 9 further including an axial compressor, wherein the atleast one fan blade further includes at least one inducer at a radiallyinward end of the compressor chamber, the at least one inducer turningsubstantially axial flow from the axial compressor toward substantiallyradial flow in the compressor chamber.
 12. The turbine engine of claim 9wherein the core airflow from the diffuser flows generally axiallyforward and then radially into a combustion chamber of the combustor.13. The turbine engine of claim 12 wherein the core airflow from thediffuser flows generally axially forward and then radially outwardlyinto the combustion chamber of the combustor.
 14. The turbine engine ofclaim 9 wherein none of the core airflow flows radially from the atleast one fan blade into the combustor.
 15. A method for operating aturbine engine including the steps of: a) centrifugally compressing coreairflow within each of a plurality of fan blades rotating in a plane; b)turning at least substantially all of the compressed core airflow out ofthe plane toward a combustion chamber, wherein none of the core airflowflows radially from the fan blades into the combustion chamber; and c)mixing the compressed core airflow with fuel only after the compressedcore airflow is turned out of the plane in said step b).
 16. The methodof claim 15 wherein the plane does not intersect the combustion chamber.17. The method of claim 15 wherein all of the compressed core airflow isturned out of the plane toward the combustion chamber in said step b).18. The method of claim 15 wherein at least substantially all of thecompressed core airflow is turned axially forward out of the planetoward the combustion chamber in said step b).